The present disclosure relates to gas turbine engines, and in particular, to a fan case for a gas turbine engine.
The fan section of a gas turbine engine includes an array of fan blades which project radially from a hub within a fan case. Although exceedingly unlikely, it is possible for a fan blade or a fragment thereof to separate from the hub and strike the fan case. The fan case operates to prevent any liberated material from radially exiting the engine. The demands of blade containment are balanced by the demands for low weight and high strength.
For relatively small diameter engines, adequate containment capability is typically achieved with a hardwall design in which a metallic case thick enough to resist penetration by a blade fragment is utilized. For relatively large diameter engines, a metallic fan case thick enough to resist penetration may be prohibitively heavy so a softwall design is typically utilized in which a light weight, high strength ballistic fabric is wrapped in a plurality of layers around a relatively thin, penetration susceptible metallic or composite case. In operation, a separated blade fragment penetrates the case and strikes the fabric. The case is punctured locally but retains structural integrity after impact. The punctured case continues to support the fabric and maintain clearance for the blade tips.
In turbofan engines, differences between the fan blade material and fan case material may contribute to thermally induced rub. Turbine engine fans and their cases experience differential thermal expansion across an operational range. For example, in flight, where other portions of the engine are subject to heating, the fan and fan case temperatures may decrease at altitude. An exemplary temperature decrease from ground to altitude may be in excess of 120 F (50 C). With an exemplary metallic fan blades and non metallic fan case, the decrease in temperature may cause the fan to decrease in diameter more than the fan case due to the coefficient of thermal expansion differential.